A gas turbine engine generally includes a fan and a core arranged in flow communication with one another. Additionally, the core of the gas turbine engine generally includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air is provided from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to atmosphere.
More specifically, the combustion section includes a combustor having a combustion chamber defined by a combustor liner. Downstream of the combustor, the turbine section includes one or more stages, for example, each stage may contain a plurality of stationary nozzle airfoils arranged in a turbine nozzle assembly, as well as a plurality of blade airfoils attached to a rotor that is driven by the flow of combustion gases against the blade airfoils. The turbine section may have other configurations as well. In any event, a typical turbine nozzle assembly includes an inner boundary, generally referred to as an inner band, and an outer boundary, generally referred to as an outer band, and the plurality of nozzle airfoils extend from the inner band to the outer band.
The aft ends of typical turbine nozzle inner and outer bands, i.e., the portion of the inner and outer bands near a trailing edge of each nozzle airfoil, experience high temperatures that distress and degrade the durability of the bands. For instance, holes for a film of cooling air must be defined upstream of a throat between airfoils to minimize an aerodynamic efficiency loss. As a result, most of the cooling film dissipates before reaching the aft region of the bands such that the cooling film largely does not benefit the aft region of the inner and outer bands. As another example, a thermal barrier coating (TBC) applied to the inner and outer bands is prone to spallation, which reduces the effectiveness of the TBC, and the TBC typically has a limited effect in areas where cooling of the opposite surface is minimal, e.g., in the aft regions in the inner and outer turbine nozzle bands.
Accordingly, improved turbine nozzle aft band portions would be desirable. For example, a turbine nozzle assembly utilizing a high temperature material, such as a ceramic matrix composite (CMC) material, in the aft region of the turbine nozzle inner band and the aft region of the turbine nozzle outer band would be beneficial. As a particular example, a turbine nozzle system comprising a turbine nozzle segment with axially cropped inner and outer bands, a CMC inner member forming an aft portion of the inner band, and a CMC outer member forming an aft portion of the outer band, would be advantageous. Additionally, a turbine nozzle system having a CMC shroud comprising the CMC outer member would be useful.